The present disclosure relates to a tapered thermal barrier coating applied to surfaces of a turbine engine component, such as a turbine blade.
As turbine inlet temperatures increase to improve engine thrust and cycle efficiency, advanced technologies are needed to cool the trailing edges of turbine blades while minimizing the amount of cooling flow used. The use of refractory metal cores to create high density patterns of cast cooling features generally provides high convective heat transfer at low cooling flow requirements. In turbine airfoil applications, the thermal heat load at the trailing edge of the airfoil is higher on the pressure side, or concave, airfoil surface relative to the suction side, or convex, airfoil surface.
Center discharge cooling circuits have been formed using a variety of fabrication techniques including, but not limited to, using refractory metal core. Such cooling is desirable to assist in the reduction of metal temperatures and to help achieve turbine life goals. Despite such a cooling circuit, there remains a large thermal gradient from the pressure side to the suction side due to the mismatch of heat loads. This thermal mismatch may increase the thermal strain across the airfoil, and may result in low thermal-mechanical fatigue life.